Many supersonic aircraft employ gas turbine engines that are capable of propelling the aircraft at supersonic speeds. These gas turbine engines, however, generally operate on subsonic flow in a range of about Mach 0.3 to 0.6 at the upstream face of the engine. The inlet decelerates the incoming airflow to a speed compatible with the requirements of the gas turbine engine. To accomplish this, a supersonic inlet is comprised of a compression surface and corresponding flow path, used to decelerate the supersonic flow into a strong terminal shock. Downstream of the terminal shock, subsonic flow is further decelerated using a subsonic diffuser to a speed corresponding with requirements of the gas turbine engine.
As is known in the art, the efficiency of the supersonic inlet and the diffusion process is a function of how much total pressure is lost in the air stream between the entrance side of the inlet and the discharge side. The total-pressure recovery of an inlet is defined by a ratio of total pressure at the discharge to total pressure at freestream.
Supersonic inlets are typically either “2D”, having a rectangular opening, or axisymmetric, having a circular opening. The supersonic inlet includes a throat positioned between a converging supersonic diffuser and a diverging subsonic diffuser. Supersonic inlets are generally also classified into three types: internal compression, mixed compression, and external compression.
Internal compression inlets accomplish supersonic and subsonic compression completely within the interior of the inlet duct. The primary theoretical advantage of this inlet type is the extremely low cowling angle that results from a completely internalized shock train. While this inlet design appears theoretically advantageous, in practice it requires a complex and performance-penalizing shock control system in order to position the shock train, to “start” the inlet, and to maintain dynamic shock stability to avoid the inlet's high sensitivity to shock train expulsion (“unstart”). The challenges associated with this type of inlet have limited its use to primarily air-breathing missile applications designed for high Mach number. Below speeds of about Mach 3.5, mixed compression and external compression inlets offer a more practical compromise between performance and complexity.
As the name implies, mixed compression inlets offer a blending of external and internal compression and seek a more practical balance between performance and complexity than that offered by fully internal compression designs in the Mach range from approximately 2.5 to 3.5. The internal portion of the shock train of a mixed compression inlet is less sensitive to flow disturbances than a fully internal design, and has lower cowling angle and drag than a fully external compression inlet designed to the same speed. But mixed compression nevertheless requires a complex control system for starting the internal shock train and for stability management to avoid inlet unstart. Two notable applications of mixed compression include the inlets on the XB-70 Valkyrie and SR-71 Blackbird aircraft.
External compression inlets are most appropriate for applications below about Mach 2.5. In this speed range, external compression offers a design simplicity that typically outweighs its generally inferior pressure recovery. Because the shock train is completely external, cowling angles, and therefore installed drag characteristics, tend to be higher when compared against internal and mixed compression designs at similar speed. However, because the shock train on an external compression inlet remains completely outside of the internal flow path, it is not subject to the sudden unstart expulsion produced by upstream or downstream flow disturbances. External compression shock stability is therefore superior to mixed or internal compression designs, requiring a significantly less complicated inlet control system. Notable examples of inlets employing external compression include those used on the Concorde, the F-14 Tomcat, and the F-15 Eagle.
Traditional inlet design methods have generally focused on improving propulsion system performance by maximizing total inlet pressure recovery and hence gross engine thrust. Complicated secondary systems and variable geometry inlets are often used to accomplish this. While high pressure recovery definitely provides certain gains, maximizing pressure recovery typically comes at the price of significant inlet drag and inlet complexity, characteristics that typically run counter to a robust and low cost-of-operation design.
For example, attempts to increase pressure recovery include bleed air-based methods, which, as is understood in the art, improve inlet pressure recovery through shock strength management and boundary layer removal. The Concorde used a method of bleed air extraction at the inlet throat that weakened the strength of the terminal shock thereby improving total pressure recovery. However, bleed air-based methods typically take a large portion of the intake flow to produce the desired results and suffer corresponding drag-related penalties once the bleed flow is eventually dumped back overboard. Additionally, extensive secondary systems are typically required, consisting of complex flow routing equipment.
Inlet ramp positioning is another method used to improve pressure recovery through more optimum placement of the compression shock system, particularly at off-design operating conditions. The Concorde, F-14, and F-15 are all examples of aircraft that employ ramp positioning for improved pressure recovery. However, ramp positioning requires electric or hydraulic actuators and an inlet control system, resulting in a large increase in inlet part count and complexity. Such systems introduce potential failure points and add significantly to development and operating costs.
The traditional supersonic inlet design process begins with the selection of compression surface geometry that best meets the performance and integration requirements of the intended application, for example aircraft design speed and/or terminal shock Mach number. For an external compression inlet, a compression surface configuration typically focuses the inlet-generated shocks, at supersonic design cruise speed, at a location immediately forward of the cowl highlight or cowl lip, generally referred to as shock-on-lip focusing. This arrangement generally provides good pressure recovery, low flow spillage drag, and a predictable post-shock subsonic flow environment that lends itself to more basic analytical techniques and explains the technique's traceability to the earliest days of supersonic inlet design.
External compression inlet design practice also uses cowl lip angle to align the cowling lip with the local supersonic flow in the vicinity of the terminal shock and the cowl lip. Aligning the lip with the local flow helps to prevent the formation of an adverse subsonic diffuser flow area profile or a complex internal shock structure in the lip region, which reduce inlet pressure recovery and flow pumping efficiency, as well as undermine diffuser flow stability.
However, as understood in the art, as supersonic design speed increases, so does the amount of compression necessary to decelerate the flow to a fixed terminal shock Mach number. Additional compression implies the need for more flow-turning off of the inlet axis, resulting in a corresponding increase in the cowl lip angle (in order to align the cowl lip angle with the local flow at the terminal shock). Any increase in cowl lip angle results in additional inlet frontal area, increasing inlet drag as speed increases. This adverse trend is a key reason why conventional external compression inlets lose viability at high supersonic Mach numbers.
One attempt to control cowl lip drag, as discussed in U.S. Pat. No. 6,793,175 issued to Sanders, includes configuring the inlet to minimize the shape and size of the cowl. Sanders' concept involves morphing a traditional rectangular intake into a more complex, but higher performance, 3-D geometry that, in a frontal view, initially resembles a circumferential sector of an axisymmetric intake, but now with the compression surface on the outer radius and the cowling on the inner radius. The cowl side extends across a similar circumferential angular arc in a frontal view, but because is it located on an inner radius, the physical arc of the cowl is reduced. The cowl drag is said to be effectively lessened through a reduction in transcribed circumferential distance. The practicality of this inlet concept is reduced by aircraft integration challenges created by the 3-D geometry. For example, the cross-sectional shape may be more difficult to integrate from a packaging perspective than an equivalent axisymmetric design for podded propulsion systems. In addition, the complex inlet shape is likely to create complex distortion patterns that require either large scale mitigating techniques in the subsonic diffuser or the use of engines with more robust operability characteristics.
Another method to reduce cowl lip angle to reduce drag involves decreasing the flow turn angle by increasing the inlet terminal shock Mach number. However, the improvement in installed drag in using a higher terminal shock Mach number is often offset by the loss in thrust from the reduction in pressure recovery resulting from the stronger terminal shock. As understood by those in the art, increasing the terminal shock Mach number also encounters significant limitations in practice once viscous flow effects are introduced. Higher terminal shock Mach numbers aggravate the shock-boundary layer interaction and reduce shock base boundary layer health. The increase in shock strength in the base region also reduces inlet buzz margin, reducing subcritical flow throttling capability. Additionally, the increase in terminal shock Mach number ultimately increases the likelihood for the need of a complex boundary layer management or inlet control system
Inlet compression surfaces are typically grouped as either ‘straight’ or ‘isentropic.’ An isentropic surface generally represents a continuously curved surface that produces a continuum of infinitesimally weak shocklets during the compression process. By contrast, a straight surface generally represents flat ramp or conic sections that produce discrete oblique or conic shocks. While an inlet employing an isentropic surface can have theoretically better pressure recovery than an inlet employing a straight-surface designed to the same operating conditions, real viscous effects combine to reduce the overall performance of isentropic inlets and can lead to poorer boundary layer health when compared to their equivalent straight-surface counterparts. Both straight and isentropic inlet types conventionally designed to the same terminal shock Mach number also produce similar flow turn angle at the cowl lip and, consequently, similar cowl lip angles. As such, neither the straight-surface inlet design nor the conventional isentropic inlet design provides a cowl drag benefit relative to the other.
As such, conventional design provides no significant latitude for adjusting the geometric arrangement of inlet and the cowl lip when designing a mechanically simple inlet compression surface using conventional shock-on-lip focusing. Because the isolated cowl drag characteristics are relatively inflexible, inlet drag relief has historically been limited to minimizing inlet-airframe interference effects.